Gas turbine engines, such as those utilized in military and commercial aircraft, include a compressor section that compresses air from an inlet and provides the compressed air to a combustor. The combustor mixes the compressed air with a fuel and ignites the mixture. The resultant gasses from the ignition are expelled into a turbine section, and expansion of the gasses through the turbine section causes the turbine section to rotate.
Each of the compressor, combustor and turbine sections are fluidly connected by a primary flow path through which air and combustion gasses travel axially through the engine. Airflow characteristics through the flow path are generally controlled via static vanes and other engine structures mounted in the flow path. The vanes and other structures are disposed circumferentially about the engine.
Gas turbine engines of the above described type frequently include cooling systems that require the transmission of cooling air, or other coolants, from one section of the gas turbine engine to another. When this transmission is required to pass through the core flow path, the coolant is passed through a pass through passage in a vane or other static structure. When the fluid in the core flow path at the static structure containing the pass through is excessively hot, heat can transfer through the vane or other static structure into the coolant passing through the pass through passage. This heat transfer can require the coolant to be overcooled prior to being passed through the vane or other static structure in order to ensure that the coolant being provided on the other side of the flow path is not too hot. Alternatively, other mitigation means are used to ensure that the coolant entering the vane or other static structure is sufficiently cool that any heat picked up by the coolant while passing through the vane or other static structure does not render the coolant too hot.